RU2383474C1 - Method and device to control mutliengined aircraft thrust - Google Patents

Abstract

FIELD: engines and pumps.

SUBSTANCE: proposed method consists in determining the parametre that defines current weight to be derived from said parametre, calculating reduced thrust proportional to aforesaid current weight and generating instruction to be sent to aircraft healthy engines. The latter generate reduced thrust complying with calculated one. The invention covers also the device to this end.

EFFECT: improved flight performances with no reduction in payload.

7 cl, 2 dwg

Description

This invention relates to a method and apparatus for controlling the thrust of an aircraft, in particular a transport aircraft, multi-engine and, in particular, twin-engine type.

It is known that usually on an airplane the vertical stabilizer (i.e., the stationary part of the vertical tail of the aircraft, designed to ensure its navigational stability) and the rudder (i.e., a movable flap mounted on a vertical stabilizer that can be maneuvered to change the direction of the aircraft), allowing the aircraft to fully comply with various flight requirements (one way or another described in the rules), in particular, such as the minimum evolving speed (VMC speed b) in case of failure of one of the engines.

It is also known that to improve the flight characteristics of an aircraft (such as fuel consumption, noise level, etc.) without reducing the payload, designers have to reduce the weight of the aircraft, if possible, namely to reduce the weight of the structure itself, its components, equipment, etc. .d., and in particular, the mass of the mentioned vertical stabilizer and rudder. However, reducing the size of the vertical stabilizer and rudder for such purposes automatically increases the aforementioned VMC speed to the detriment of the flight characteristics of the aircraft because if one of the engines fails, the minimum speed that must be achieved increases.

The present invention relates to a method for controlling the thrust of a multi-engine aircraft, which eliminates the above disadvantages.

To this end, in accordance with this invention, said method is distinguished by the following:

a) determine a parameter that characterizes the current weight of the aircraft, and said current weight is derived from this parameter. This parameter can be, for example, the current weight of the aircraft itself, or the angle of attack;

b) calculate the reduced thrust value, which is proportional to the current weight determined in step a);

c) determine at least one command per engine, a command that is such that, if it is applied to those engines of the aircraft that have not failed, these engines produce a thrust substantially equal to the reduced thrust calculated in step b); and

d) the command referred to in step c) is applied to those aircraft engines that have not failed.

The method in accordance with the present invention can be applied to an aircraft in which all engines are running. However, in a preferred embodiment of this method, the aircraft engines should be monitored so that it is possible to detect a failure of one of the aforementioned engines, and steps a) and d) are carried out only if one of the engines is found to fail.

So, on the basis of the invention, when one of the aircraft engines fails, the thrust of the engine or engines that did not fail decreases (as a function of the weight of the aircraft) so that the angle of inclination also decreases, but at the same time remains steep enough stable, preventing undesirable positions of the aircraft in flight, if the aircraft is forced to enter the second circle when its mass is small. As a result of decreasing the yaw moment of this or these engines or engines that have not failed, theoretically the minimum evolutionary speed of the VMC decreases regardless of the weight of the aircraft. Thus, it becomes possible to balance the yawing aircraft without casting a rudder. This means that the speed will not be limited by the minimum speed of the VMC. Thus, restrictions on flight characteristics, due to the minimum evolutionary speed due to the failure of one of the engines, are less stringent, and the maximum benefit from the flight characteristics of the longitudinal movement of the aircraft can be achieved by reducing the flight speed, if necessary.

As a result, an aircraft that implements the method in accordance with this invention significantly reduces flight performance limitations due to the minimum VMC speed. The various aforementioned flight performance requirements can thus be achieved regardless of the weight of the aircraft, regardless of its speed. This, in particular, is particularly advantageous for a twin-engine aircraft. This is due to the fact that in a twin-engine aircraft, the thrust of one engine should provide a control minimum angle of inclination with a large mass. The result of this is that, with a lower mass, the thrust of an individual engine is greater than the thrust necessary to achieve a standard tilt angle.

Moreover, the method in accordance with this invention also allows you to align the angle and spatial position of the aircraft, especially with all running engines.

In a preferred embodiment of the method in step b), the reduced thrust value FOEI is calculated using the following expression:

FOEI = mg (γcons + 1 / f),

* where mg is the mentioned current weight, m is mass, and g is the acceleration of gravity;

* γcons is the value of the angle of inclination corresponding to the control angle of inclination; but

* f is the current relationship between lift and drag.

In such a preferred embodiment, said inclination value γcons is calculated using the following expression:

γcons = γmin + Δγ,

Where

* γmin is the normative value of the minimum value of the angle of inclination of the aircraft; and

* Δγ is a predefined value of the angle of inclination, for example, several degrees.

In addition, in a specific embodiment, in step a), the current weight of the aircraft is determined by summing the weight of the aircraft without fuel and the (current) weight of the fuel on board the aircraft.

In this case, it is also effective during the approach phase to perform the following actions:

- to step d):

* the first value of the weight of the aircraft is determined by summing the weight of the aircraft without fuel and the weight of the fuel on board the aircraft;

* the second weight value is determined as a function of the current angle of attack of the aircraft;

* the mentioned first and second weight values are compared with each other; and

- in step d):

* if the aforementioned first and second values are equal or are within an acceptable limit, then the said command determined in step c) is applied to those engines that did not fail to obtain a thrust substantially equal to the reduced thrust value calculated in step b ); and

* if the first and second weight values differ (by more than the permissible limit), the command is sent to those engines that did not fail, in order to get a constant predetermined draft, for example a type of TOGA (where TOGA is a well-known abbreviation of the English expression “Take Off” / Go Around ”(“ To take off / Go to the second round ”).

Thus, when the current weight (namely, the first weight value referred to), which is determined taking into account the weight of the fuel on board the aircraft, and which is used to determine the aforementioned reduced thrust value, differs from the (second value) weight calculated as a function of the current angle of attack, a reduced value of traction according to this invention is not applied. This is due to the fact that, in this case, the weight value obtained through the angle of attack is usually determined very accurately. This means that the current weight value determined in step a) is probably erroneous and thus leads to the erroneous value of the reduced traction force calculated during step b), and is also erroneous for the command determined in step c) method in accordance with this invention.

The present invention also relates to a thrust control device for a multi-engine aircraft, in particular a multi-engine transport aircraft and, in particular, a twin-engine aircraft.

In accordance with the invention, said device of this type preferably includes means for monitoring the engines of said aircraft in order to be able to detect a failure of one of said engines, characterized in that it further includes:

- the first means for determining a parameter characterizing the current weight of the aircraft, and for outputting said current weight from it;

- second means for calculating a reduced thrust value, which is proportional to said current weight determined by said first means;

- third means for determining at least one engine command, which is such that, if it is applied to those engines that have not failed, these engines produce a thrust substantially equal to said reduced thrust value calculated by said second means; and

- the fourth means for giving a command, defined by the third means, to those engines of the aircraft that did not fail.

The device in accordance with this invention, therefore, makes it possible, in the event of a failure of one of the engines, to adapt the thrust of this or these engines (engines), which did not fail or did not fail, so as to obtain the desired flight characteristics, but with a lower minimum evolving speed (VMC speed) and a more stable position of the aircraft in flight.

The figures of the attached diagrams make it easy to understand how the device will be implemented. In these figures, the same footnotes mean the same elements.

Figure 1 is a structural diagram of a control device in accordance with this invention.

Figure 2 schematically shows a specific improvement of the control device in accordance with this invention.

The device 1 in accordance with this invention, schematically depicted in figure 1, is designed to control the thrust of a multi-engine aircraft, in particular multi-engine transport aircraft, not shown in figure 1.

Although, and not only, the aforementioned device 1 is applicable to a greater extent to a twin-engine aircraft, and it is preferable to use it at least during the approach phase for landing on the runway.

In addition, in particular, said device 1 is intended for use when one of the engines (not shown) of a multi-engine aircraft fails. To this end, said device 1 comprises conventional means 2 designed to control aircraft engines in such a way as to be able to detect a failure of one of said engines.

According to the invention, the device 1 further includes the following:

- means 3 for determining a parameter representing the current weight of the aircraft, and for outputting said current weight from it. This parameter can be either the current weight itself or the angle of attack, for example, it will be possible to determine the current weight using the usual lifting force equation;

- means 4 for calculating a reduced thrust value, which is proportional to the current weight determined by said means 3;

- means 5, which is connected to the means 4 through the data channel 4A and which is designed in such a way as to determine at least one command for the engine. This command is such that when it is applied to one or those engine or engines that have not failed, this or these engine (s) will produce a thrust substantially equal to the reduced thrust calculated by means 4. In one particular embodiment of means 4 and 5 form part of a central control unit (UC), which, for example, is connected via data channel 6 to said means 3 and via data channel 13 to means 2; and

- conventional means 7, which is connected via a data transmission channel 8 to a central control unit (UC) and which is formed in such a way as to give the aforementioned command, determined by means 5, to that (those) engine (s) that did not fail (and). This conventional means 7 may, for example, provide a command to an autonomous digital engine control system of the FADEC type (FADEC is the generally recognized acronym of the English expression “Full Authority Digital Engine Control” - “Autonomous digital engine control system”).

Thus, when one of the aircraft engines fails, the device 1 in accordance with this invention reduces the thrust (as a function of the weight of the aircraft) of one or those engines (engines) that remain operational, in order to reduce the angle of inclination of the aircraft while at the same time providing a sufficient angle of inclination to perform maneuvers. As a result, by reducing the yaw moment of one or those engine (s) that have not failed, the theoretical minimum evolutionary speed (VMC speed) is reduced regardless of the weight of the aircraft. Thus, it becomes possible to balance the yawing aircraft without casting a rudder. This means that the speed will not be limited by the minimum VMC speed. Thus, the limitations on flight performance due to the minimum evolving speed due to failure of one of the engines become less stringent, and the maximum benefit from the flight performance of the longitudinal movement of the aircraft can be achieved by reducing the flight speed, if necessary.

As a result, an aircraft using device 1 in accordance with this invention significantly reduces flight performance limitations due to minimum VMC speed. Different flight requirements can thus be achieved regardless of the weight of the aircraft and regardless of its speed. This is especially advantageous for a twin engine aircraft.

Moreover, the device 1 in accordance with this invention also has the advantage that it can be used even when all the engines of the aircraft are in order to achieve a more stable spatial position of the aircraft, to protect the engines and thus reduce damage. Plus, it saves on maintenance costs.

In a preferred embodiment of the method, means 4 calculates a reduced thrust FOEI value using the following relationship:

FOEI = mg (γcons + 1 / f), where

* mg - current weight, m - mass, and g - acceleration of gravity;

* γcons is the value of the angle of inclination equal to the control angle of inclination; but

* f is the current relationship between lift and drag.

In a preferred embodiment, means 4 also calculates the value of the angle of inclination from the following expression:

γcons = γmin + Δγ,

Where

* γmin - standard value of the minimum angle of inclination of the aircraft; but

* Δγ is a predefined value of the angle of inclination, for example, 0.5 ° degrees.

This device 1 also includes a series of information sources 9, which are connected by a connection 10 to the central control unit UC and which are capable of supplying to the control unit values of a plurality of parameters, in particular, the values of the aforementioned current ratio f.

In this embodiment, the device 1 further includes a control system (SC), shown schematically in figure 2. This control system SC is used when the landing site is in line of sight, in particular:

- means 14, for example, similar to said means 3 or corresponding to means 3, which is formed in such a way as to determine the first value of the weight of the aircraft. For this, said means 14 (as well as means 3) summarize the weight of the aircraft without fuel with the weight of the fuel on board the aircraft. Usually, the weight of the aircraft without fuel is determined by the pilot and entered into the device 1 using a conventional input means 15, which, as a rule, is a multifunctional control and display unit type MCDU (where MCDU is the common abbreviation for the English expression “Multifunction Control Display Unit ”-“ Multifunctional control and display unit ”), which is connected through a data channel 16 with said means 14;

- the means 17 in the usual way determines the second value of the weight of the aircraft as a function of the current angle of attack, which, for example, is obtained from the usual means 18, the latter is connected via the data channel 19 to the said means 17. This means 18 can be part of the said set of information sources 9. In addition, said means 17 comprises a conventional circuit used to determine the second weight value from the lift equation based, for example, on the measurement of the angle of attack during flight. Mass can be determined from this equation by correlating it with speed;

- means 20, which is connected via data channels 21 and 22 with means 14 and 17 and which is intended to compare the first and second weight values with each other and generate a command depending on this comparison.

In accordance with this invention, the purpose of said means 20 is to prohibit the application of the aforementioned reduced thrust value to that (those) engine (s) that have not failed when said first and second weight values are different from each other. To this end, said means 20 may, for example, transmit an inhibit command to the central control unit UC via the data channel 23.

Thus, based on the SC control system:

- if the first and second weight values are equal (or are within an acceptable limit, for example, equal to several percent of the current mass), then device 1 sends the above command to that (those) engine (s) that did not fail, in order to obtain traction, substantially equal to the reduced thrust value that is calculated by means 4; and

- if the first and second weight values differ from each other (by more than a predetermined value), then device 1 gives the usual command to that (those) engine (s) that did not fail to obtain a constant predetermined traction, for example, TOGA type traction force (where TOGA is the generally accepted abbreviation of the English expression “Take Off / Go Around” (“Take off / Go to the second round”).

As a result, if the current weight (i.e., the first weight value), which is determined by means 14 taking into account the weight of the fuel on board the aircraft and which is used to determine the reduced thrust, is different from the weight calculated by means 17 as a function of the current angle attack, the device 1 in accordance with this invention does not use the obtained reduced thrust value, since during the approach (during which the said control system is preferably activated) value the weight obtained from the angle of attack is usually very accurate, which means that the current weight determined by means 14 and 3 is probably erroneous. The use of this erroneous value leads to an incorrect determination of the reduced value of the thrust force calculated by means 4 and given by the command determined by the apparatus 5.

Claims (7)

1. A method of controlling the thrust of a multi-engine aircraft, in this method: c) determine at least one engine command, a command that is such that if it is supplied to those aircraft engines that have not failed, these engines issue thrust substantially equal to the reduced thrust value, and d) said command determined in step c) is applied to those engines of the aircraft that have not failed, in which, in order to determine said reduced thrust value used in step c ): a) a parameter characterizing the current weight of the aircraft is determined, and said current weight is derived from this parameter, and b) as a reduced value of the thrust force, a reduced thrust value FOEI is calculated, which is proportional to the said current weight of the aircraft, this is done using the following expression : FOEI = mg (γcons + 1 / f), where mg is the mentioned current weight, m is mass, ag is the acceleration of gravity; γcons is the value of the angle of inclination corresponding to the reference angle of inclination and f is the current ratio between the lifting force and the drag of the aircraft. 2. The method according to claim 1, in which, while the aircraft is in flight, the engines of said aircraft are controlled in such a way as to be able to detect a failure of one of said engines, and steps a) and d) are performed only if when a failure of one of the mentioned engines is detected. 3. The method according to claim 1, in which the said value of the angle of inclination γcons is calculated using the following expression: γcons = γmin + Δγ,
where γmin is the normative value of the minimum angle of inclination for the aircraft, and Δγ is the predetermined value of the angle of inclination. 4. The method according to claim 1, in which at step a) the current value of the weight of the aircraft is determined by summing the weight of the aircraft without fuel with the weight of the fuel on board the aircraft. 5. The method according to claim 1, in which during the approach phase, the following operations are also performed: before step d) the first value of the weight of the aircraft is determined by summing the weight of the aircraft without fuel with the weight of the fuel on board the aircraft, the second value weights are determined as a function of the current angle of attack of the aircraft, the first and second weight values are compared with each other, and in step d), if the aforementioned first and second weight values are equal to the allowable limit, then for engines Those who did not fail, give the said command in order to obtain a thrust essentially equal to the reduced thrust value calculated in step b), and if the first and second weight values are different, then those engines that did not fail give a command in order to get a constant predetermined traction. 6. A device for controlling the thrust of a multi-engine aircraft, said device comprising means (5) for determining at least one engine command, such that if it is supplied to those engines that have not failed, these engines give thrust, essentially equal to the reduced thrust value, and means (7) for supplying the aforementioned command to those aircraft engines that did not fail, means (3) for determining a parameter characterizing the current weight of the aircraft, and for outputting from said current weight, and means (4) for calculating, as a reduced thrust value, a reduced thrust value FOEI, which is proportional to said current weight, using the following expression: FOEI = mg (γcons + 1 / f), where mg is said current weight , m is the mass, ag is the acceleration of gravity, γcons is the value of the angle of inclination corresponding to the control angle of inclination, af is the current ratio between the lifting force and the drag of the aircraft. 7. An aircraft containing a device (1) similar to that described in claim 6.

Images